For aircraft approach operations in mountainous regions, flight paths where strict adherence to a narrow corridor is imposed when the conditions of visibility are poor or military navigation at very low altitudes (sometimes referred to as flight in the terrain), the performance characteristics of the system of navigation must meet a level of demand compatible with the needs of the non-precision approaches commonly denoted by the acronym RNP (for “Required Navigation Precision”). The level of demand varies depending on the phase of the flight or the mission set. The level of demand consists in meeting criteria for precision, integrity, continuity and availability. The demand on integrity corresponds to the capacity of the system of navigation to raise an alarm within a given time delay when the system is affected by a failure or by a fault leading to a risk of error on the measurement of positioning, greater than a certain threshold.
In order to quantify the integrity of a position measurement, a parameter called protection radius is used. The protection radius corresponds to a maximum position error for a given level of confidence (or of risk). In other words, the probability of the position error exceeding the stated protection radius without an alarm being sent to a navigation system is less than this given probability value. The given probability value is commonly called the non-integrity level or level of risk. It is expressed as risk per hour that the position error will exceed the calculated protection radius without an alarm being raised. For operations of the RNP type in the approach phase, the non-integrity level generally required goes from 10−7 to 10−9 occurrences of a position error, greater than the protection radius, per hour depending on the type of failure envisaged.
A satellite positioning receiver commonly called GNSS (for “Global Navigation Satellite System”) supplies information on position and speed of the carrier by triangulation using the signals emitted by orbiting satellites visible from the carrier. For civilian or non-classified military applications, the GNSS receiver also supplies the raw pseudo-range and Doppler measurements corresponding to each satellite tracked. In some civilian applications, the satellite positioning receiver is included in a more global multi-channel radiofrequency receiver system called MMR (for “Multi-Mode receiver”) notably comprising an ILS receiver dedicated to the landing phases. “Satellite positioning receiver” is therefore understood to mean all the parts of the multi-channel radiofrequency receiver (multi-mode receiver or MMR) involved in the reception of the raw measurements.
The geographical positioning information supplied by the GNSS receiver may be momentarily unavailable because the receiver must have a minimum of four satellites of the positioning system in direct view in order to be able to generate a measurement. Furthermore, the satellite data may contain errors due to failures affecting the satellites. These unreliable data values must then be identified so as not to invalidate the position coming from the GNSS receiver or from an INS/GNSS hybrid position (described hereinbelow) which would use these raw measurements. In order to detect satellite failures and ensure the integrity of the GNSS measurements, a known solution is to equip a satellite positioning receiver with a precision and availability estimation system known as a RAIM (for “Receiver Autonomous Integrity Monitoring”) which is based on the geometry and the redundancy of the constellation of satellites in order to exclude the faulty satellite.
More recently, hybrid systems have been developed which mathematically combine the inertial measurements supplied by an inertial measurement unit IMU and the measurements supplied by the satellite positioning receiver GNSS. These hybrid systems are known in the prior art by the term INS/GNSS (“Inertial Navigation System” and “Global Navigation Satellite System”). An inertial navigation system is defined which is a device including an inertial measurement unit IMU and associated processing means (notably enabling pure inertial navigation and inertia/GPS hybridization calculations to be performed). An inertial navigation system is also called an IRU (for “Inertial Reference Unit”). An inertial navigation system also including a processor or a partition of a processor designed to calculate flight parameters with respect to the air (air speed, incidence angle, wind sheer, altitude, etc.) and connected to probes situated on the skin of the aircraft which measure the total pressure, the static pressure, the temperature, etc. is called reference inertial and aerodynamic navigation system and ADIRU (for “Air Data Inertial Reference Unit”).
The calculations for pure inertia navigation, based on the use of the measurements from the IMU (gyrometers and accelerometers) provide accurate position information in the short term, but which drift over the long term (under the influence of the defects in the sensors). In contrast, the measurements supplied by a GNSS receiver are less accurate in the short term, but do not drift over time. Thus, the precision of the errors supplied by the GNSS receiver allow the inertial drift to be controlled and the low-noise inertial measurements allow the noise on the measurements of the GNSS receiver to be filtered. The combination of the two types of measurement is carried out by a Kalman filter.
Kalman filtering is based, on the one hand, on the possibilities for modelling the time variation of the state of a physical system considered in its environment, by means of an equation referred to as trend equation (pre-estimation), and for modelling of the dependence relationship existing between the states of the physical system in question. On the other hand, the measurements by which the physical system is perceived from the outside are modelled, by means of an equation referred to as observation equation, in order to allow the re-adjustment of the states of the filter (post-estimation), in other words to enable the Kalman filter to be re-adjusted. In a Kalman filter, the effective measurement or “measurement vector” allows an estimate to be made of the state of the system retroactively which is optimal in the sense that it minimizes the covariance of the error made on this estimation. The estimator part of the filter generates pre-estimates of the vector of state of the system by using the observed difference between the effective measurement vector and its theoretical prediction to generate a corrective term, called ‘innovation’.
One exemplary method for determination of positions and for monitoring their integrity is described in the Patent application WO2008040658 filed by the applicant. In this method, the hybridization is carried out using a Kalman filter. The Kalman filter receives the position and speed points supplied by the inertial measurement unit and the positioning measurements (pseudo-range and delta-range) supplied by the satellite positioning unit (known as ‘tight hybridization’ or ‘on satellite axes’), models the time variation of the errors of the inertial navigation system and delivers the estimate of these errors retroactively. This method is known as a ‘closed-loop method’ since the estimate of the errors is used retroactively to readjust the positioning and speed point of the inertial navigation system. The method could also be carried out in open-loop mode, but the post-estimate of the inertial errors would then be sub-optimal with respect to the closed-loop process, since the estimation of a part of these variables would no longer perfectly correspond to the assumption of linearity.
In the document WO2008040658, a method for monitoring the integrity of the position measurement continually calculates a protection radius associated with the measurement of the position of the object in flight. As described in the document WO2008040658, the evaluation of the value of the protection radius generally results from calculations of probabilities using the static precision characteristics of the GNSS measurements and of the behaviour of the inertial sensors. The calculation is based on two types of error which are, on the one hand, the normal measurement errors and, on the other, the errors caused by an operational anomaly of the constellation of satellites. Furthermore, they take into account the eventuality of an undetected satellite failure.
However, the devices for positioning and for monitoring the integrity of the position measurement, that are based on hybridization methods or on systems for estimating precision and availability known as RAIM, do not guarantee the position measurements in the case of hardware failures in the on-board positioning receivers in the aircraft that remain undetected (by the equipment itself), namely the GNSS receivers or the inertial navigation systems. In other words, the value of the protection radius does not take into account the eventuality, and even less the presence, of a hardware failure affecting one of the positioning receivers. A hardware failure not detected by the built-in tests internal to the receivers and to the inertial navigation systems is then capable of generating an error, on the calculated position, greater than the aforementioned protection radius (which only takes into account satellite failures) without an alarm being raised. Unfortunately, the frequency of occurrence of an undetected hardware failure in the on-board equipment is often greater than the level of integrity required for the non precision approach phases. It is therefore necessary to protect the calculated position from their effect.